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  1. FIFTY YEARS OF FLIGHT RESEARCH: AN ANNOTATED BIBLIOGRAPHY OF TECHNICAL PUBLICATIONS OF NASA DRYDEN FLIGHT RESEARCH CENTER, 1946-1996 , Technical Publication
    Authors: David F. Fisher
    Report Number: NASA-TP-1999-206568
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Titles, authors, report numbers, and abstracts are given for more than 2200 unclassified and unrestricted technical reports and papers published from September 1946 to December 1996 by NASA Dryden Flight Research Center and its predecessor organizations. These technical reports and papers describe and give the results of 50 years of flight research performed by the NACA and NASA, from the X-1 and other early X-airplanes, to the X-15, Space Shuttle, X-29 Forward Swept Wing, and X-31 aircraft. Some of the other research airplanes tested were the D-558, phase 1 and 2; M-2, HL-10 and X-24 lifting bodies; Digital Fly-By-Wire and Supercritical Wing F-8; XB-70; YF-12; AFTI F-111 TACT and MAW; F-15 HiDEC; F-18 High Alpha Research Vehicle, and F-18 Systems Research Aircraft. The citations of reports and papers are listed in chronological order, with author and aircraft indices. In addition, in the appendices, citations of 233 contractor reports, more than 200 UCLA Flight System Research Center reports and 25 video tapes are included.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 99
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          Postscript (9,609 KBytes)
          PDF (9,968 KBytes)
    Report Date: May 1999
    No. Pages: 525
    Funding Organization: WU 953 36 00 000 RR 00 000
    Keywords:      Bibliographies; Databases; Flight-tests; High-speed flight aircraft; Research facilities; Research vehicles
    Notes: Unclassified and unlimited reports from 1946 through 1996.


  2. FLIGHT-DETERMINED, SUBSONIC, LATERAL-DIRECTIONAL STABILITY AND CONTROL DERIVATIVES OF THE THRUST-VECTORING F-18 HIGH ANGLE OF ATTACK RESEARCH VEHICLE (HARV), AND COMPARISONS TO THE BASIC F-18 AND PREDICTED DERIVATIVES , Technical Paper
    Authors: Kenneth W. Iliff and Kon-Sheng Charles Wang
    Report Number: NASA-TP-1999-206573
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The subsonic, lateral-directional, stability and control derivatives of the thrust-vectoring F-18 High Angle of Attack Research Vehicle (HARV) are extracted from flight data using a maximum likelihood parameter identification technique. State noise is accounted for in the identification formulation and is used to model the uncommanded forcing functions caused by unsteady aerodynamics. Preprogrammed maneuvers provided independent control surface inputs, eliminating problems of identifiability related to correlations between the aircraft controls and states. The HARV derivatives are plotted as functions of angles of attack between 10 deg and 70 deg and compared to flight estimates from the basic F-18 aircraft and to predictions from ground and wind-tunnel tests. Unlike maneuvers of the basic F-18 aircraft, the HARV maneuvers were very precise and repeatable, resulting in tightly clustered estimates with small uncertainty levels. Significant differences were found between flight and prediction; however, some of these differences may be attributed to differences in the range of sideslip or input amplitude over which a given derivative was evaluated, and to differences between the HARV external configuration and that of the basic F-18 aircraft, upon which most of the prediction was based. Some HARV derivative fairings have been adjusted using basic F-18 derivatives (with low uncertainties) to help account for differences in variable ranges and the lack of HARV maneuvers at certain angles of attack.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 08
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    Report Date: January 1999
    No. Pages: 90
    Funding Organization: WU 529-50-04-00-RR-00-000
    Keywords:      Aerodynamics; F-18 aircraft; HARV; High angle of attack; Maximum likelihood estimation; Parameter estimation; Parameter identification; Stability and control derivatives; Thrust vectoring; Wind-tunnel predictions
    Notes: Kon-Sheng Charles Wang is a member of the technical staff of The Aerospace Corporation, P.O. Box 92957, Los Angeles, CA 90009-2957.


  3. A REASSESSMENT OF HEAVY-DUTY TRUCK AERODYNAMIC DESIGN FEATURES AND PRIORITIES , Technical Paper
    Authors: Edwin J. Saltzman and Robert R. Meyer, Jr.
    Report Number: NASA-TP-1999-206574
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Between 1973 and 1982, the NASA Dryden Flight Research Center conducted 'coast-down' tests demonstrating means for reducing the drag of trucks, buses, and motor homes. Numerous configurations were evaluated using a box-shaped test van, a two-axle truck, and a tractor-semitrailer combination. Results from three configurations of the test van are of interest now in view of a trucking industry goal of a 0.25 drag coefficient for tractor-semitrailer combinations. Two test van configurations with blunt-base geometry, similar to present day trucks (one configuration has square front corners and the other has rounded front corners), quantify the base drag increase associated with reduced forebody drag. Hoerner's equations predict this trend; however, test van results, reinforced by large-scale air vehicle data, indicate that Hoerner's formula greatly underestimates this dependence of base drag on forebody efficiency. The demonstrated increase in base drag associated with forebody refinement indicates that the goal of a 0.25 drag coefficient will not be achieved without also reducing afterbody drag. A third configuration of the test van had a truncated boattail to reduce afterbody drag and achieved a drag coefficient of 0.242. These results are included here and references are identified for other means of reducing afterbody drag.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02, 31, 37, 85, 99
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    Report Date: June 1999
    No. Pages: 38
    Funding Organization: WU 251-10-01-00-TT-00-000
    Notes: Edwin J. Saltzman, Analytical Services & Materials, Edwards, California; Robert R. Meyer, Jr., Dryden Flight Research Center, Edwards, California.


  4. FLIGHT TEST OF AN ADAPTIVE CONFIGURATION OPTIMIZATION SYSTEM FOR TRANSPORT AIRCRAFT , Technical Memorandum
    Authors: Glenn B. Gilyard, Jennifer Georgie and Joseph S. Barnicki
    Report Number: NASA-TM-1999-206569
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A NASA Dryden Flight Research Center program explores the practical application of real-time adaptive configuration optimization for enhanced transport performance on an L-1011 aircraft. This approach is based on calculation of incremental drag from forced-response, symmetric, outboard aileron maneuvers. In real-time operation, the symmetric outboard aileron deflection is directly optimized, and the horizontal stabilator and angle of attack are indirectly optimized. A flight experiment has been conducted from an onboard research engineering test station, and flight research results are presented herein. The optimization system has demonstrated the capability of determining the minimum drag configuration of the aircraft in real time. The drag-minimization algorithm is capable of identifying drag to approximately a one-drag-count level. Optimizing the symmetric outboard aileron position realizes a drag reduction of 2-3 drag counts (approximately 1 percent). Algorithm analysis of maneuvers indicate that two-sided raised-cosine maneuvers improve definition of the symmetric outboard aileron drag effect, thereby improving analysis results and consistency. Ramp maneuvers provide a more even distribution of data collection as a function of excitation deflection than raised-cosine maneuvers provide. A commercial operational system would require airdata calculations and normal output of current inertial navigation systems; engine pressure ratio measurements would be optional.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02, 03
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          Postscript (308 KBytes)
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    Report Date: January 1999
    No. Pages: 17
    Funding Organization: WU 522-16-14-00-39-00-L10
    Keywords:      Adaptive control; Ailerons; Aircraft performance; Cambered wings; Commercial aircraft; Drag reduction; Flight optimization; Fuel consumption; Optimization; Transport aircraft; Flaps (control surfaces); Variable geometry; Wing camber
    Notes: Presented at 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 11-14, 1999,AIAA-99-0831.


  5. A PARALLEL, FINITE-VOLUME ALGORITHM FOR LARGE-EDDY SIMULATION OF TURBULENT FLOWS , Technical Memorandum
    Authors: Trong T. Bui
    Report Number: NASA-TM-1999-206570
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A parallel, finite-volume algorithm has been developed for large-eddy simulation (LES) of compressible turbulent flows. This algorithm includes piecewise linear least-square reconstruction, trilinear finite-element interpolation, Roe flux-difference splitting, and second-order MacCormack time marching. Parallel implementation is done using the message-passing programming model. In this paper, the numerical algorithm is described. To validate the numerical method for turbulence simulation, LES of fully developed turbulent flow in a square duct is performed for a Reynolds number of 320 based on the average friction velocity and the hydraulic diameter of the duct. Direct numerical simulation (DNS) results are available for this test case, and the accuracy of this algorithm for turbulence simulations can be ascertained by comparing the LES solutions with the DNS results. The effects of grid resolution, upwind numerical dissipation, and subgrid-scale dissipation on the accuracy of the LES are examined. Comparison with DNS results shows that the standard Roe flux-difference splitting dissipation adversely affects the accuracy of the turbulence simulation. For accurate turbulence simulations, only 3-5 percent of the standard Roe flux-difference splitting dissipation is needed.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 34
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    Report Date: January 1999
    No. Pages: 26
    Funding Organization: WU 529-50-04
    Keywords:      Finite-volume; Large-eddy simulation; Parallel computers; Turbulence; Upwindmethods
    Notes: Presented at 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 11-14, 1999,AIAA-99-0789.


  6. WIND AND MOUNTAIN WAVE OBSERVATIONS FOR THE PATHFINDER HAWAIIAN FLIGHT TEST OPERATION , Contractor Report
    Authors: Edward H. Teets, Jr. and Natalie Salazar
    Report Number: NASA-CR-1999-206571
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Light-wing-loaded aircraft such as the high altitude solar electric powered aircraft Pathfinder (designed by AeroVironment, Inc., Simi Valley, Ca.) are extremely sensitive to the wind speeds, wind shears, and turbulence and thereby are easily affected by mountain gravity wave activity. After reaching an altitude over 67,400 ft on June 9, 1997 the Pathfinder aircraft encountered a mountain wave during descent produced by the 5,000+ feet mountain of Hawaii's northernmost island of Kauai. This paper will discuss the atmospheric conditions and aircraft configuration for this case of mountain wave observed at the U. S. Navy's Pacific Missile Range Facility (PMRF) at Barking Sands, Kauai, Hawaii. This paper also describes the pathfinder airplane, climatology, and atmospheric conditions leading to the wave formation.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 47
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    Report Date: January 1999
    No. Pages: 9
    Funding Organization: WU 529-10-04-00-29-00-ERA
    Keywords:      Atmospheric Waves; Flight test; Meteorology; Mountain waves; Solar-powered
    Notes: Presented at 8th American Meterological Society Conference on Aviation Range & Aerospace Meterology, Jan. 10-15, 1999, Dallas, TX. E. Teets, AS&M, Inc., Edwards, CA; N. Salazar, New Mexico Highlands Univ., NM. NASA Technical Monitor: L.J. Ehernberger.


  7. FLIGHT-DETERMINED SUBSONIC LIFT AND DRAG CHARACTERISTICS OF SEVEN LIFTING-BODY AND WING-BODY REENTRY VEHICLE CONFIGURATIONS WITH TRUNCATED BASES , Conference Paper
    Authors: Edwin J. Saltzman, K. Charles Wang and Kenneth W. Iliff
    Report Number: H-2287
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper examines flight-measured subsonic lift and drag characteristics of seven lifting-body and wing-body reentry vehicle configurations with truncated bases. The seven vehicles are the full-scale M2-F1, M2-F2, HL-10, X-24A, X-24B, and X-15 vehicles and the Space Shuttle prototype. Lift and drag data of the various vehicles are assembled under aerodynamic performance parameters and presented in several analytical and graphical formats. These formats unify the data and allow a greater understanding than studying the vehicles individually allows. Lift-curve slope data are studied with respect to aspect ratio and related to generic wind-tunnel model data and to theory for low- aspect-ratio planforms. The proper definition of reference area was critical for understanding and comparing the lift data. The drag components studied include minimum drag coefficient, lift-related drag, maximum lift-to-drag ratio, and, where available, base pressure coefficients. The effects of fineness ratio on forebody drag were also considered. The influence of forebody drag on afterbody (base) drag at low lift is shown to be related to Hoerner's compilation for body, airfoil, nacelle, and canopy drag. These analyses are intended to provide a useful analytical framework with which to compare and evaluate new vehicle configurations of the same generic family.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02 and 15
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    Report Date: January 1999
    No. Pages: 32
    Funding Organization: 529-50-04-T2-RR-00-000
    Keywords:      Aerodynamics; Lifting bodies; Reentry vehicles; Reusable launch vehicles
    Notes: Presented as AIAA 99-0383 at the 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 11-14, 1999.


  8. FACTORS AFFECTING INLET-ENGINE COMPATIBILITY DURING AIRCRAFT DEPARTURES AT HIGH ANGLE OF ATTACK FOR AN F/A-18A AIRCRAFT , Technical Memorandum
    Authors: W. G. Steenken, J. G. Williams, A. J. Yuhas and K. R. Walsh
    Report Number: NASA-TM-1999-206572
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The F404-GE-400 engine powered F/A-18A High Alpha Research Vehicle (HARV) was used to examine the quality of inlet airflow during departed flight maneuvers, that is, during flight outside the normal maneuvering envelope where control surfaces have little or no effectiveness. A series of six nose-left and six nose-right departures were initiated at Mach numbers between 0.3 and 0.4 at an altitude of 35 kft. The yaw rates at departure recovery were in the range of 40 to 90 degrees per second. Engine surges were encountered during three of the nose-left and one of the nose-right departures. Time-variant inlet-total-pressure distortion levels at the engine face were determined to not significantly exceed those measured at maximum angle-of-attack and -sideslip maneuvers during controlled flight. Surges as a result of inlet distortion levels were anticipated to initiate in the fan. Analysis revealed that the surges initiated in the compressor and were the result of a combination of high levels of inlet distortion and rapid changes in aircraft motion. These rapid changes in aircraft motion are indicative of a combination of engine mount and gyroscopic loads being applied to the engine structure that impact the aerodynamic stability of the compressor through changes in the rotor-to-case clearances.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 07
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    Report Date: February 1999
    No. Pages: 198
    Funding Organization: WU 529-31-04-00-37-00-F18
    Keywords:      Aircraft departures; F-18 aircraft; F/A-18A; High angle of attack; Inlet distortion; Inlet-engine compatibility
    Notes: W. G. Steenken and J. G. Williams, GE Aircraft Engines, Cincinnati, Ohio; A. J. Yuhas, Analytical Services and Materials, Inc., Edwards, California; and K. R. Walsh, Dryden Flight Research Center, Edwards, California. NASA Contract NAS3-26617.


  9. A BASE DRAG REDUCTION EXPERIMENT ON THE X-33 LINEAR AEROSPIKE SR-71 EXPERIMENT (LASRE) FLIGHT PROGRAM , Conference Report
    Authors: Stephen A. Whitmore and Timothy R. Moes
    Report Number: H-2328
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Drag reduction tests were conducted on the LASRE/X-33 flight experiment. The LASRE experiment is a flight test of a roughly 20-percent scale model of an X-33 forebody with a single aerospike engine at the rear. The experiment apparatus is mounted on top of an SR-71 aircraft. This paper suggests a method for reducing base drag by adding surface roughness along the forebody. Calculations show a potential for base drag reductions of 8 to 14 percent. Flight results corroborate the base drag reduction, with actual reductions of 15 percent in the high-subsonic flight regime. An unexpected result of this experiment is that drag benefits were shown to persist well into the supersonic flight regime. Flight results show no overall net drag reduction. Applied surface roughness causes forebody pressures to rise and offset base drag reductions. Apparently the grit displaced streamlines outward, causing forebody compression. Results of the LASRE drag experiments are inconclusive and more work is needed. Clearly, however, the forebody grit application works as a viable drag reduction tool.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: n.a.
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          Postscript (937 KBytes)
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    Report Date: January 1999
    No. Pages: 17
    Funding Organization: 242-33-02-00-23-00-T15
    Keywords:      n.a.
    Notes: Presented at the 37th AIAA Aerospace Sciences meeting and Exhibit, January 11-14, 1999, Reno, Nevada. AIAA 99-0277.


  10. THE HISTORY OF SOLID-PROPELLANT ROCKETRY: WHAT WE DO AND DO NOT KNOW , Conference Paper
    Authors: J. D. Hunley
    Report Number: H-2330
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Contributions to the evolution of solid-propellant rocketry have come from a variety of sources. World War II research on large solids enabled one company to capitalize on work in the area of castable double-base propellants. Separate development of castable composite propellants led to production of Polaris and Minuteman powerplants. Pivotal to the development of these missiles were Edward Hall's advocacy of the Minuteman missile within the Air Force and contract funding to resolve problems. The discovery that adding large amounts of aluminum significantly increased the specific impulse of a castable composite propellant further aided large-missile technology. These separate lines of research led to the development of large solid-propellant motors and boosters. Many more discoveries went into the development of large solid-propellant motors. Ammonium perchlorate replaced potassium perchlorate as an oxidizer in the late 1940's, and binders were developed. Discoveries important in the evolution of large solid-propellant motors appear to have resulted from innovators' education and skills, an exposure to contemporary problems, an awareness of theory but a willingness not to let it dictate empirical investigations, and proper empirical techniques. Other important contributions are the adequate funding and exchange of information. However, many questions remain about these and other innovations.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 15 and 28
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    Report Date: June 1999
    No. Pages: 12
    Funding Organization: WU 038-00-00-L0-XR-00-000
    Keywords:      Solid propellants; Solid rocket booster; Space launch vehicles; Space shuttle
    Notes: Presented as AIAA 99-2925 Invited Paper at the 35th AIAA, ASME, SAE, ASEE Joint Propulsion Conference and Exhibit, Los Angeles, California, June 20-24, 1999.


  11. SIMULATOR EVALUATION OF SIMPLIFIED PROPULSION-ONLY EMERGENCY FLIGHT CONTROL SYSTEMS ON TRANSPORT AIRCRAFT , Technical Memorandum
    Authors: Frank W. Burcham, Jr., Trindel A. Maine, John Kaneshige and John Bull
    Report Number: NASA-TM-1999-206578
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: With the advent of digital engine control systems, considering the use of engine thrust for emergency flight control has become feasible. Many incidents have occurred in which engine thrust supplemented or replaced normal aircraft flight controls. In most of these cases, a crash has resulted, and more than 1100 lives have been lost. The NASA Dryden Flight Research Center has developed a propulsion-controlled aircraft (PCA) system in which computer-controlled engine thrust provides emergency flight control capability. Using this PCA system, an F-15 and an MD-11 airplane have been landed without using any flight controls. In simulations, C-17, B-757, and B-747 PCA systems have also been evaluated successfully. These tests used full-authority digital electronic control systems on the engines. Developing simpler PCA systems that can operate without full-authority engine control, thus allowing PCA technology to be installed on less capable airplanes or at lower cost, is also a desire. Studies have examined simplified 'PCA Ultralite' concepts in which thrust control is provided using an autothrottle system supplemented by manual differential throttle control. Some of these concepts have worked well. The PCA Ultralite study results are presented for simulation tests of MD-11, B-757, C-17, and B-747 aircraft.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 08
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    Report Date: June 1999
    No. Pages: 49
    Keywords:      B-747; B-757; C-17; Emergency flight control; MD-11; Propulsive control.
    Notes: Frank Burcham, Jr., Trindel Maine, NASA Dryden, Edwards, CA; John Kaneshige, NASA Ames, Moffett Field, CA; and John Bull, CAELUM Research Corporation, NASA Ames, Moffett Field, CA. WU 522-35-14-00-33-00-IDA


  12. OPEN-MODE DEBONDING ANALYSIS OF CURVED SANDWICH PANELS SUBJECTED TO HEATING AND CRYOGENIC COOLING ON OPPOSITE FACES , Technical Paper
    Authors: William L. Ko
    Report Number: NASA-TP-1999-206580
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Increasing use of curved sandwich panels as aerospace structure components makes it vital to fully understand their thermostructural behavior and identify key factors affecting the open-mode debonding failure. Open-mode debonding analysis is performed on a family of curved honeycomb-core sandwich panels with different radii of curvature. The curved sandwich panels are either simply supported or clamped, and are subjected to uniform heating on the convex side and uniform cryogenic cooling on the concave side. The finite-element method was used to study the effects of panel curvature and boundary condition on the open-mode stress (radial tensile stress) and displacement fields in the curved sandwich panels. The critical stress point, where potential debonding failure could initiate, was found to be at the midspan (or outer span) of the inner bonding interface between the sandwich core and face sheet on the concave side, depending on the boundary condition and panel curvature. Open-mode stress increases with increasing panel curvature, reaching a maximum value at certain high curvature, and then decreases slightly as the panel curvature continues to increase and approach that of quarter circle. Changing the boundary condition from simply supported to clamped reduces the magnitudes of open-mode stresses and the associated sandwich core depth stretching.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 39
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    Report Date: June 1999
    No. Pages: 43
    Funding Organization: WU 242-33-02-00-23-00-TA1
    Keywords:      Curved sandwich panels; Debonding stresses; Open-mode debonding; Thermo-cryogenic loading


  13. A BASE DRAG REDUCTION EXPERIMENT ON THE X-33 LINEAR AEROSPIKE SR-71 EXPERIMENT (LASRE) FLIGHT PROGRAM , Technical Memorandum
    Authors: Stephen A. Whitmore and Timothy R. Moes
    Report Number: NASA-TM-1999-206575
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Drag reduction tests were conducted on the LASRE/X-33 flight experiment. The LASRE experiment is a flight test of a roughly 20-percent scale model of an X-33 forebody with a single aerospike engine at the rear. The experiment apparatus is mounted on top of an SR-71 aircraft. This paper suggests a method for reducing base drag by adding surface roughness along the forebody. Calculations show a potential for base drag reductions of 8 to 14 percent. Flight results corroborate the base drag reduction, with actual reductions of 15 percent in the high-subsonic flight regime. An unexpected result of this experiment is that drag benefits were shown to persist well into the supersonic flight regime. Flight results show no overall net drag reduction. Applied surface roughness causes forebody pressures to rise and offset base drag reductions. Apparently the grit displaced streamlines outward, causing forebody compression. Results of the LASRE drag experiments are inconclusive and more work is needed. Clearly, however, the forebody grit application works as a viable drag reduction tool.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: n.a.
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          Postscript (825 KBytes)
          PDF (327 KBytes)
    Report Date: January 1999
    No. Pages: 21
    Funding Organization: WU 242-33-02-00-23-00-T15
    Keywords:      Aerospike engine; Base drag; Skin friction.
    Notes: Presented at the 37th AIAA Aerospace Sciences meeting and Exhibit, January 11-14, 1999, Reno, Nevada. AIAA 99-0277.


  14. UNSTEADY AERODYNAMICS - SUBSONIC COMPRESSIBLE INVISCID CASE
    Authors: A. V. Balakrishnan
    Report Number: NASA-CR-1999-206583
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper presents a new analytical treatment of Unsteady Aerodynamics - the linear theory covering the subsonic compressible (inviscid) case - drawing on some recent work in Operator Theory and Functional Analysis. The specific new results are: (a) An existence and uniqueness proof for the Laplace transform version of the Possio integral equation as well as a new closed form solution approximation thereof. (b) A new representation for the time-domain solution of the subsonic compressible aerodynamic equations emphasizing in particular the role of the initial conditions.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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    Report Date: August 1999
    No. Pages: 81
    Keywords:      Boundary value problem; Closed form solution; Field equations; Initial value problem;Laplace transform; Operator theory; Possio equation; Semigroups; Subsonic compressible inviscid aerodynamics; Unsteady aerodynamics.
    Notes: Kenneth W. Iliff submitted this report on disks, then through 5 PDF Files for distribution of this contractor report to CASI and others. NASA DFRC grant NCC2-374 with UCLA.


  15. SELECTED PERFORMANCE MEASUREMENTS OF THE F-15 ACTIVE AXISYM METRIC THRUST-VECTORING NOZZLE
    Authors: John S. Orme and Robert L. Sims
    Report Number: H-2339
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight tests recently completed at the NASA Dryden Flight Research Center evaluated performance of a hydromechanically vectored axisymmetric nozzle onboard the F-15 ACTIVE. A flight-test technique whereby strain gages installed onto engine mounts provided for the direct measurement of thrust and vector forces has proven to be extremely valuable. Flow turning and thrust efficiency, as well as nozzle static pressure distributions were measured and analyzed. This report presents results from testing at an altitude of 30,000 ft and a speed of Mach 0.9. Flow turning and thrust efficiency were found to be significantly different than predicted, and moreover, varied substantially with power setting and pitch vector angle. Results of an in-flight comparison of the direct thrust measurement technique and an engine simulation fell within the expected uncertainty bands. Overall nozzle performance at this flight condition demonstrated the F100-PW-229 thrust-vectoring nozzles to be highly capable and efficient.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 05 and 07
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    Report Date: September 1999
    No. Pages: 12
    Keywords:      Axisymmetric nozzle; Direct thrust measurement; F-15 ACTIVE; Flight tests; Nozzle aerodynamic loads; Nozzle drag; Nozzle flow field; Nozzle performance; Thrust efficiency; Thrust measurement; Thrust vectoring
    Notes: Presented at the 14th ISABE (International Society for Airbreathing Engines) Annual Symposium, September 5-10, 1999, Florence, Italy. ISABE Paper No. IS 166.


  16. DESIGN AND PREDICTIONS FOR A HIGH-ALTITUDE (LOW-REYNOLDS-NUMBER) AERODYNAMIC FLIGHT EXPERIMENT , Technical Memorandum
    Authors: Donald Greer, Phil Hamory, Keith Krake and Mark Drela
    Report Number: NASA-TM-1999-206579
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A sailplane being developed at NASA Dryden Flight Research Center will support a high-altitude flight experiment. The experiment will measure the performance parameters of an airfoil at high altitudes (70,000 to 100,000 ft), low Reynolds numbers (200,000 to 700,000), and high subsonic Mach numbers (0.5 and 0.65). The airfoil section lift and drag are determined from pitot and static pressure measurements. The locations of the separation bubble, Tollmien-Schlichting boundary layer instability frequencies, and vortex shedding are measured from a hot-film strip. The details of the planned flight experiment are presented. Several predictions of the airfoil performance are also presented. Mark Drela from the Massachusetts Institute of Technology designed the APEX-16 airfoil, using the MSES code. Two-dimensional Navier-Stokes analyses were performed by Mahidhar Tatineni and Xiaolin Zhong from the University of California, Los Angeles, and by the authors at NASA Dryden.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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          Postscript (1,505 KBytes)
          PDF (461 KBytes)
    Report Date: July 1999
    No. Pages: 23
    Funding Organization: 529-1004-0029-00-APE
    Keywords:      Airfoils; High altitude; Low Reynolds numbers; Sailplane; Transition
    Notes: Presented at the 17th Applied Aerodynamics Conference and 14th Computational Fluid Dynamics Conference, Norfolk, Virginia, June 28-July 1, 1999, AIAA 99-3183. This information was also published in the Journal of Aircraft, Vol. 37, No. 4, July-August 2000.


  17. INITIAL FLIGHT TEST OF THE PRODUCTION SUPPORT FLIGHT CONTROL COMPUTERS AT NASA DRYDEN FLIGHT RESEARCH CENTER
    Authors: John Carter and Mark Stephenson
    Report Number: NASA-TM-1999-206581
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The NASA Dryden Flight Research Center has completed the initial flight test of a modified set of F/A-18 flight control computers that gives the aircraft a research control law capability. The production support flight control computers (PSFCC) provide an increased capability for flight research in the control law, handling qualities, and flight systems areas. The PSFCC feature a research flight control processor that is 'piggybacked' onto the baseline F/A-18 flight control system. This research processor allows for pilot selection of research control law operation in flight. To validate flight operation, a replication of a standard F/A-18 control law was programmed into the research processor and flight-tested over a limited envelope. This paper provides a brief description of the system, summarizes the initial flight test of the PSFCC, and describes future experiments for the PSFCC.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 08
    Availability:
        Format(s) on-line:
          Postscript (1,226 KBytes)
          PDF (516 KBytes)
    Report Date: August 1999
    No. Pages: 19
    Keywords:      Aircraft testing; Control law research; F/A-18; Flight control; Production support
    Notes: Presented as AIAA 99-4203 at the AIAA Guidance, Navigation, and Control Conference, Portland, Oregon, August 9-11, 1999.


  18. CRYOGENIC FUEL TANK DRAINING ANALYSIS MODEL , Conference Paper
    Authors: Donald Greer
    Report Number: H-2344
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: One of the technological challenges in designing advanced hypersonic aircraft and the next generation of spacecraft is developing reusable flight-weight cryogenic fuel tanks. As an aid in the design and analysis of these cryogenic tanks, a computational fluid dynamics (CFD) model has been developed specifically for the analysis of flow in a cryogenic fuel tank. This model employs the full set of Navier-Stokes equations, except that viscous dissipation is neglected in the energy equation. An explicit finite difference technique in two-dimensional generalized coordinates, approximated to second-order accuracy in both space and time is used. The stiffness resulting from the low Mach number is resolved by using artificial compressibility. The model simulates the transient, two-dimensional draining of a fuel tank cross section. To calculate the slosh wave dynamics the interface between the ullage gas and liquid fuel is modeled as a free surface. Then, experimental data for free convection inside a horizontal cylinder are compared with model results. Finally, cryogenic tank draining calculations are performed with three different wall heat fluxes to demonstrate the effect of wall heat flux on the internal tank flow field.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 34
    Availability:
        Format(s) on-line:
          Postscript (1,959 KBytes)
          PDF (510 KBytes)
    Report Date: July 1999
    No. Pages: 9
    Funding Organization: WU 519-10-44-E8-RI-00-000
    Keywords:      Computation fluid dynamics; Cryogenic; Free surface; Fuel tank; Numerical modeling.
    Notes: Presented as paper CLB-3 at the CEC/ICMC 1999 Cryogenic Engineering and International Cryogenic Materials Conference, July 12-16, 1999, Montreal, Quebec, Canada.


  19. RECONFIGURABLE FLIGHT CONTROL DESIGNS WITH APPLICATION TO THE X-33 VEHICLE
    Authors: John J. Burken, Ping Lu and Zhenglu Wu
    Report Number: NASA-TM-1999-206582
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Two methods for control system reconfiguration have been investigated. The first method is a robust servomechanism control approach (optimal tracking problem) that is a generalization of the classical proportional-plus-integral control to multiple input-multiple output systems. The second method is a control-allocation approach based on a quadratic programming formulation. A globally convergent fixed-point iteration algorithm has been developed to make onboard implementation of this method feasible. These methods have been applied to reconfigurable entry flight control design for the X-33 vehicle. Examples presented demonstrate simultaneous tracking of angle-of-attack and roll angle commands during failures of the right body flap actuator. Although simulations demonstrate success of the first method in most cases, the control-allocation method appears to provide uniformly better performance in all cases.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 08
    Availability:
        Format(s) on-line:
          Postscript (511 KBytes)
          PDF (197 KBytes)
    Report Date: August 1999
    No. Pages: 20
    Keywords:      Failure control system design; Reconfigurable flight controls; Robust
    Notes: Presented at AIAA Guidance Navigation and Control Conference, Portland, Oregon, August 9-11, 1999, AIAA-99-4134.


  20. DESIGN AND CALIBRATION OF AN AIRBORNE MULTICHANNEL SWEPT-TUNED SPECTRUM ANALYZER , Technical Memorandum
    Authors: Philip J. Hamory, John K. Diamond and Arild Bertelrud
    Report Number: NASA-TM-1999-206584
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper describes the design and calibration of a four-channel, airborne, swept- tuned spectrum analyzer used in two hypersonic flight experiments for characterizing dynamic data up to 25 kHz. Built mainly from commercially available analog function modules, the analyzer proved useful for an application with limited telemetry bandwidth, physical weight and volume, and electrical power. The authors discuss considerations that affect the frequency and amplitude calibrations, limitations of the design, and example flight data.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 06
    Availability:
        Format(s) on-line:
          Postscript (270 KBytes)
          PDF (177 KBytes)
    Report Date: October 1999
    No. Pages: 17
    Keywords:      Spectrum analysis; Signal processing; Multiple input calibration;Multivariable calibration; Analog circuits
    Notes: Presented at the 35th Annual International Telmetering Conference (ITC) 'Telemetry: Meeting The 21st Century Challenge,' Las Vegas, Nevada, October 25-28, 1999.


  21. 1998 RESEARCH ENGINEERING ANNUAL REPORT , Technical Memorandum
    Authors: Compiled by Gerald N. Malcolm
    Report Number: NASA-TM-1999-206585
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Selected research and technology activities at Dryden Flight Research Center are summarized. These activities exemplify the Center's varied and productive research efforts.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 99
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          Postscript (20,147 KBytes)
          PDF (6,311 KBytes)
    Report Date: August 1999
    No. Pages: 78
    Funding Organization: WU 953-36-00-GH-RR-00-000
    Keywords:      Aerodynamics; Flight; Flight controls; Flight systems; Flight test; Instrumentation;
    Notes: Point of Contact: Contact person(s) at the end of each article.


  22. ESTIMATED BENEFITS OF VARIABLE-GEOMETRY WING CAMBER CONTROL FOR TRANSPORT AIRCRAFT.
    Authors: Alexander Bolonkin and Glenn B. Gilyard
    Report Number: NASA-TM-1999-206586
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Analytical benefits of variable-camber capability on subsonic transport aircraft are explored. Using aerodynamic performance models, including drag as a function of deflection angle for control surfaces of interest, optimal performance benefits of variable camber are calculated. Results demonstrate that if all wing trailing-edge surfaces are available for optimization, drag can be significantly reduced at most points within the flight envelope. The optimization approach developed and illustrated for flight uses variable camber for optimization of aerodynamic efficiency (maximizing the lift-to-drag ratio). Most transport aircraft have significant latent capability in this area. Wing camber control that can affect performance optimization for transport aircraft includes symmetric use of ailerons and flaps. In this paper, drag characteristics for aileron and flap deflections are computed based on analytical and wind-tunnel data. All calculations based on predictions for the subject aircraft and the optimal surface deflection are obtained by simple interpolation for given conditions. An algorithm is also presented for computation of optimal surface deflection for given conditions. Benefits of variable camber for a transport configuration using a simple trailing-edge control surface system can approach more than 10 percent, especially for nonstandard flight conditions. In the cruise regime, the benefit is 1-3 percent.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02-01, 02-03, 03-01
    Availability:
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          Postscript (628 KBytes)
          PDF (370 KBytes)
    Report Date: October 1999
    No. Pages: 49
    Keywords:      Adaptive control; Aircraft performance; Cambered wings; Commercial aircraft;
    Notes: Alexander Bolonkin, Senior Research Associate of the National Research Council, Washington D. C. and Glenn B. Gilyard, NASA Dryden Flight Research Center, Edwards, California.


  23. INLET FLOW CHARACTERISTICS DURING RAPID MANEUVERS FOR AN F/A-18A AIRCRAFT , Technical Memorandum
    Authors: William G. Steenken, John G. Williams and Kevin R. Walsh
    Report Number: NASA-TM-1999-206587
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The F404-GE-400 engine powered F/A-18A High Alpha Research Vehicle (HARV) was used to examine the characteristics of inlet airflow during rapid aircraft maneuvers. A study of the degree of similarity between inlet data obtained during rapid aircraft maneuvers and inlet data obtained at steady aerodynamic attitudes was conducted at the maximum engine airflow of approximately 145 lbm/sec using a computer model that was generated from inlet data obtained during steady aerodynamic maneuvers. Results show that rapid-maneuver inlet recoveries agreed very well with the recoveries obtained at equivalent stabilized angle-of-attack conditions. The peak dynamic circumferential distortion values obtained during rapid maneuvers agreed within 0.01 units of distortion over the 10 - 38 degree angle of attack range with the values obtained during steady aerodynamic maneuvers while similar agreement was found for the peak dynamic radial distortion values up to 29 degrees angle-of-attack. Exceedences of the rapid-maneuver peak dynamic circumferential distortion values relative to the peak distortion model values at steady attitudes occurred only at low or negative angles of attack and were inconsequential from an engine-stability assessment point of view. The results of this study validate the current industry practice of testing at steady aerodynamic conditions to characterize inlet recovery and peak dynamic distortion levels.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 07
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          PDF (6,948 KBytes)
    Report Date: October 1999
    No. Pages: 221
    Keywords:      Aircraft rapid maneuvers; F-18 aircraft; F/A-18A; High angle of attack; Inlet distortion; Inlet-engine compatibility
    Notes: W. G. Steenken and J. G. Williams, GE Aircraft Engines, Cincinnati, Ohio, and K. R. Walsh, Dryden Flight Research Center, Edwards, California. NASA Contract NAS3-26617.


  24. EVALUATION OF THE LINEAR AEROSPIKE SR-71 EXPERIMENT (LASRE) OXYGEN SENSOR , Technical Memorandum
    Authors: Kimberly A. Ennix, Griffin P. Corpening, Michele Jarvis and Harry R. Chiles
    Report Number: NASA-TM-1999-206589
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The Linear Aerospike SR-71 Experiment (LASRE) was a propulsion flight experiment for advanced space vehicles such as the X-33 and reusable launch vehicle. A linear aerospike rocket engine was integrated into a semi-span of an X- 33-like lifting body shape (model), and carried on top of an SR-71 aircraft at NASA Dryden Flight Research Center. Because no flight data existed for aerospike nozzles, the primary objective of the LASRE flight experiment was to evaluate flight effects on the engine performance over a range of altitudes and Mach numbers. Because it contained a large quantity of energy in the form of fuel, oxidizer, hypergolics, and gases at very high pressures, the LASRE propulsion system posed a major hazard for fire or explosion. Therefore, a propulsion-hazard mitigation system was created for LASRE that included a nitrogen purge system. Oxygen sensors were a critical part of the nitrogen purge system because they measured purge operation and effectiveness. Because the available oxygen sensors were not designed for flight testing, a laboratory study investigated oxygen- sensor characteristics and accuracy over a range of altitudes and oxygen concentrations. Laboratory test data made it possible to properly calibrate the sensors for flight. Such data also provided a more accurate error prediction than the manufacturer's specification. This predictive accuracy increased confidence in the sensor output during critical phases of the flight. This paper presents the findings of this laboratory test.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 20
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          PDF (312 KBytes)
    Report Date: November 1999
    No. Pages: 23
    Keywords:      Aerospike; Flammability limits; High speed civil transport; Linear aerospike
    Notes: Presented at the AIAA 9th International Space Planes and Hypersonic Systems Conference, Norfolk, Virginia, November 1-5, 1999.


  25. PROPELLANT FEED SYSTEM LEAK DETECTION—LESSONS LEARNED FROM THE LINEAR AEROSPIKE SR-71 EXPERIMENT (LASRE) , Technical Memorandum
    Authors: Neal Hass, Masashi Mizukami, Bradford A. Neal, Clinton St. John,, Robert J. Beil and Timothy P. Griffin
    Report Number: NASA-TM-1999-206590
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper presents pertinent results and assessment of propellant feed system leak detection as applied to the Linear Aerospike SR-71 Experiment (LASRE) program flown at the NASA Dryden Flight Research Center, Edwards, California. The LASRE was a flight test of an aerospike rocket engine using liquid oxygen and high-pressure gaseous hydrogen as propellants. The flight safety of the crew and the experiment demanded proven technologies and techniques that could detect leaks and assess the integrity of hazardous propellant feed systems. Point source detection and systematic detection were used. Point source detection was adequate for catching gross leakage from components of the propellant feed systems, but insufficient for clearing LASRE to levels of acceptability. Systematic detection, which used high-resolution instrumentation to evaluate the health of the system within a closed volume, provided a better means for assessing leak hazards. Oxygen sensors detected a leak rate of approximately 0.04 cubic inches per second of liquid oxygen. Pressure sensor data revealed speculated cryogenic boiloff through the fittings of the oxygen system, but location of the source(s) was indeterminable. Ultimately, LASRE was canceled because leak detection techniques were unable to verify that oxygen levels could be maintained below flammability limits.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 20
    Availability:
        Format(s) on-line:
          PDF (565 KBytes)
    Report Date: November 1999
    No. Pages: 38
    Keywords:      Fire safety; Hypersonic vehicles; Leak assessment; Leak detection; Rocket
    Notes: Presented at the AIAA 9th International Space Planes and Hypersonic Systems Conference, Norfolk, Virginia, November 1-5, 1999.


  26. FLUSH AIRDATA SENSING (FADS) SYSTEM CALIBRATION PROCEDURES AND RESULTS FOR BLUNT FOREBODIES , Technical Paper
    Authors: Brent R. Cobleigh , Stephen A. Whitmore , Edward A. Haering, Jr. , Jerry Borrer and V. Eric Roback
    Report Number: NASA-TP-1999-209012
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Blunt-forebody pressure data are used to study the behavior of the NASA Dryden Flight Research Center flush airdata sensing (FADS) pressure model and solution algorithm. The model relates surface pressure measurements to the airdata state. Spliced from the potential flow solution for uniform flow over a sphere and the modified Newtonian impact theory, the model was shown to apply to a wide range of blunt-forebody shapes and Mach numbers. Calibrations of a sphere, spherical cones, a Rankine half body, and the F-14, F/A-18, X-33, X-34, and X-38 configurations are shown. The three calibration parameters are well-behaved from Mach 0.25 to Mach 5.0, an angle-of-attack range extending to greater than 30& deg;, and an angle-of-sideslip range extending to greater than 15°. Contrary to the sharp calibration changes found on traditional pitot-static systems at transonic speeds, the FADS calibrations are smooth, monotonic functions of Mach number and effective angles of attack and sideslip. Because the FADS calibration is sensitive to pressure port location, detailed measurements of the actual pressure port locations on the flight vehicle are required and the wind-tunnel calibration model should have pressure ports in similar locations. The procedure for calibrating a FADS system is outlined.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 06
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          PDF (542 KBytes)
    Report Date: November 1999
    No. Pages: 32
    Funding Organization: WU 242-33-02-00-23-00-TA3
    Keywords:      Airdata calibration; Blunt forebody; FADS; Flush airdata sensing; Sphere; X-33;
    Notes: Presented at the 9th International Space Planes and Hypersonic Systems and Technologies Conference, November 1-5, 1999, Norfolk, VA, AIAA 99-4816. Brent R. Cobleigh, Stephen A. Whitmore, and Edward A. Haering, Jr., NASA Dryden, Edwards, California; Jerry Borrer, NASA Johnson, Houston, Texas; and V. Eric Roback, NASA Langley, Hampton, Virginia. A patent has been filed on this NASA invention. Equations (10)-(16) are included in this patent.


  27. AUTOMATED TESTING EXPERIENCE OF THE LINEAR AEROSPIKESR-71 EXPERIMENT (LASRE) CONTROLLER , Technical Memorandum
    Authors: Richard R. Larson
    Report Number: NASA-TM-1999-206588
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: System controllers must be fail-safe, low cost, flexible to software changes, able to output health and status words, and permit rapid retest qualification. The system controller designed and tested for the aerospike engine program was an attempt to meet these requirements. This paper describes (1) the aerospike controller design, (2) the automated simulation testing techniques, and (3) the real time monitoring data visualization structure. Controller cost was minimized by design of a single-string system that used an off-the-shelf 486 central processing unit (CPU). A linked-list architecture, with states (nodes) defined in a user-friendly state table, accomplished software changes to the controller. Proven to be fail-safe, this system reported the abort cause and automatically reverted to a safe condition for any first failure. A real time simulation and test system automated the software checkout and retest requirements. A program requirement to decode all abort causes in real time during all ground and flight tests assured the safety of flight decisions and the proper execution of mission rules. The design also included health and status words, and provided a real time analysis interpretation for all health and status data.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 05, 61, 62
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          Postscript (363 KBytes)
          PDF (341 KBytes)
    Report Date: September 1999
    No. Pages: 34
    Keywords:      Aerospike engine; Automated software testing; Controller architecture; LASRE;
    Notes: Presented at the ITEA Test and Evaluation in the Information Age Symposium, September 21-24, 1999, Atlanta, Georgia, at a poster session.


  28. BOUNDARY-LAYER TRANSITION RESULTS FROM THE F-16XL-2 SUPERSONIC LAMINAR FLOW CONTROL EXPERIMENT , Technical Memorandum
    Authors: Laurie A. Marshall
    Report Number: NASA-TM-1999-209013
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A variable-porosity suction glove has been flown on the F-16XL-2 aircraft to demonstrate the feasibility of this technology for the proposed High-Speed Civil Transport (HSCT). Boundary-layer transition data have been obtained on the titanium glove primarily at Mach 2.0 and altitudes of 53,000-55,000 ft. The objectives of this supersonic laminar flow control flight experiment have been to achieve 50- to 60-percent-chord laminar flow on a highly swept wing at supersonic speeds and to provide data to validate codes and suction design. The most successful laminar flow results have not been obtained at the glove design point (Mach 1.9 at an altitude of 50,000 ft). At Mach 2.0 and an altitude of 53,000 ft, which corresponds to a Reynolds number of 22.7 x 10 to the sixth power, optimum suction levels have allowed long runs of a minimum of 46-percent-chord laminar flow to be achieved. This paper discusses research variables that directly impact the ability to obtain laminar flow and techniques to correct for these variables.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
    Availability:
        Format(s) on-line:
          PDF (1,359 KBytes)
    Report Date: December 1999
    No. Pages: 55
    Keywords:      F-16XL; Hot film; Laminar flow control; Suction; Transition
 
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